Chambers in which liquid and gaseous propellant components or products of their gas generation interact are widely used in engineering.
A combustion chamber is known that is mainly used for gas turbines (U.S. Pat. No. 4,211,073). The chamber comprises a mixing head consisting of an internal injector face, external and middle bottoms, and a bipropellant injector oriented along the chamber axis. The chamber casing consists of an external structural envelope and in internal fire wall, between which a regenerative cooling passage is located. Side injectors are uniformly displaced along the circumference of the chamber casing. It is difficult to use this device as a combustion chamber of high-thrust liquid-propellant rocket engines (LRE), in particular because of the fact that in practice it is not always possible to reliably ensure propellant ignition when a multi-injector mixing head is used.
A liquid-propellant rocket engine (LRE) chamber is known (U.S. Pat. No. 4,621,492).
In this chamber a mixing head comprises an internal injector face, external and middle bottoms. The injectors of the mixing head are fixed in the face and bottoms. The injectors comprise three ferrules inserted concentrically one in another and located in injector bodies. The ferrules are each made with a central axial channel for feeding a liquid oxidizer. There are injectors, protruding bodies of which extend beyond the internal injector face.
This technical solution is not effective for large-size chambers, since injectors of a single type are used therein. Furthermore, the injectors are complex in design, since they consist of three ferrules concentrically inserted one in another. The known technical solution is designed for a liquid oxidizer fed along the injector axis. The design of the injector requires substantial further development for use of a gaseous oxidizer or an oxidizing gas. Besides, in some cases of large-size LRE chambers, it is not possible to consider the feeding of the igniting mixture through a central injector of the mixing head to be completely reliable and efficient.
The technical solution most similar to the LRE chamber is the LRE chamber and its composite parts used in the American SSME rocket engine (Monograph by G. G Gakhun, V. I. Baulin, V. A. Volodin et al., "Liquid-Propellant Rocket Engine Design and Engineering," Moscow, 1989, page 124, Fig. 6.40, page 135, Fig. 7.12).
The chamber comprises a casing and a mixing head. The mixing head has front, intermediate and external bottoms and injectors mounted therein. All of the injectors used in this technical solution may be functionally divided into main injectors and injectors forming oscillation-preventing partitions. The chamber casing has an external structural envelope and an internal fire wall with a ribbed external surface, between which a regenerative cooling passage is formed.
The chamber is used for LREs operating on oxygen-hydrogen propellant components. The chamber and the composite parts thereof, including ignition and mixing systems, are designed for such propellant components. The direct use of this known technical solution for LREs operating an oxygen-kerosene propellant components is impossible and requires additional inventive activity.
Different designs of LRE chamber casings are widely used in rocket engine manufacturing.
A bipropellant rocket engine is known (U.S. Pat. No. 4,894,986).
The chamber casing of this engine consists of an external structural envelope and an internal fire wall, between which a regenerative cooling passage is formed. This passage has an inlet and an outlet. The passage inlet is located on the nozzle part of the casing.
An SSME LRE chamber casing is also known (Monograph by G. G. Gakhun, V. I. Baulin, V. A. Volodin, et al., "Liquid-Propellant Rocket Engine Design and Engineering," Moscow, 1989, page 124, Fig. 6.40.5).
The casing also consists of an external structural envelope and an internal fire wall, between which a regenerative cooling passage is positioned. The cooling passage inlet is located on a nozzle of the casing.
There are possible reserves for increasing the cooling efficiency in the technical solution disclosed in "U.S. Pat. No. 4,894,986" and in the SSME engine "Monograph by G. G. Gakhun, V. I. Baulin, V. A. Volodin et al., `Liquid-Propellant Rocket Engine Design and Engineering,` Moscow, 1989, page 124, fig. 6.40.5." In the first place this is due to the fact that feeding the cooling component of the propellant in the known designs is not carried out onto the most thermally stressed places of the internal fire wall of the casing, and in the second place, due to the fact that internal cooling of the chamber through slots is not provided. Theses limitations do not make it possible to additionally increase the temperature in the external layer of combustion products, i.e. on the internal fire wall of the casing.
The technical solution most similar to the claimed chamber casing is the casing of an LRE chamber, which comprises a chamber and a nozzle which consist of an external structural envelope and an internal fire wall with a regenerative chamber cooling passage located between them, wherewith a ring-like slot of a gas-screen belt is made in the internal fire wall (U.S. Pat. No. 3,595,023).
This construction also has limitations in the part relating to reduction of the temperature of the internal fire wall of the casing in the most thermally stressed places, both during engine start and upon reaching the operational mode of the engine, and has additional reserves for increasing the efficiency of cooling the internal fire wall of the chamber casing.